The present invention generally relates to manufacturing of large-scale structures using composite materials and, more particularly, to a method of transferring large uncured composite laminates for the manufacture of large aircraft fuselage sections.
The structural performance advantages of composites, such as carbon fiber epoxy and graphite bismaleimide (BMI) materials, are widely known in the aerospace industry. Aircraft designers have been attracted to composites because of their superior stiffness, strength, and lower weight, for example. As more advanced materials and a wider variety of material forms have become available, aerospace usage of composites has increased. For the aerospace industry, composite lamination techniques that would provide faster material lay up rates, for example, would reduce the cost of manufacturing large composite structures. New and innovative composite lamination technologies are envisioned, such as the manufacture of large aircraft fuselage sections that may exceed, for example, 15 to 20 feet in diameter.
Without high speed tape laying techniques that are currently being explored, fabrication of large scale composite structures, such as commercial aircraft fuselage skins, will require the use of numerous pieces of expensive capital equipment—such as automated tape laying and fiber placement machines and large autoclaves—to achieve tape laying rates high enough to make the use of composites economical, allowing composites to replace metal as the primary structural components of the aircraft. Increasing the productivity of each machine is important to producing composite aerospace parts economically.
In the manufacture of aerospace components, current methods of maintaining the aero surface, also referred to as the outer mold line (OML), typically include laying skins into panel sections tooled to the aero surface. Small panels are assembled into larger components. Building smaller panels with equipment tooled to the aero surface requires more equipment and more processing steps than is desired.
Systems and methods for fabricating aerospace composite structures are disclosed in U.S. Pat. No. 4,693,678, issued Sep. 15, 1987, U.S. Pat. No. 4,780,262, issued Oct. 25, 1988, U.S. Pat. No. 6,012,883, issued Jan. 11, 2000, U.S. Pat. No. 6,168,358, issued Jan. 2, 2001, and U.S. Pat. No. 5,746,553, issued May 5, 1998, which are incorporated by reference.
Curing of composite laminate parts after lay up generally entails placing the composite laminate part into an autoclave. Curing is usually performed in a pressure environment inside an autoclave, for example, to eliminate voids within the material that could be caused by air bubbles, and that could cause serious degradation of the strength of the material in the final product. A compaction force on the skin laminate is commonly provided by “bagging” the panel, in which a plastic membrane is placed over the panel and sealed to the tool. Vacuum is then applied, inside the sealed plastic “bag”, to the part while pressure is added to the autoclave and outside the bag, causing a differential pressure that compacts the laminate. In that manner, compaction pressures of greater than one atmosphere can be achieved. For composite material placed on the inside of a panel section tool, the aero surface side of the panel remains against the tool, and the non-aero, inside surface of the panel faces the vacuum bag. For a barrel section lay-up on the outside of a mandrel tool, the inside surface of the barrel section remains against the tool and the outside, aero surface of the panel faces the vacuum bag.
If one were to lay up and cure, for example, a fuselage barrel section on a male mandrel, the bag side surface would also be the aerodynamic surface of the barrel section. Since the bag side surface typically is aerodynamically rough and is difficult to make aerodynamically smooth, the result does not meet aero-smoothness requirements. Therefore, if one were to make a large barrel section, one would need to transfer it. That is, to maximize the efficiency of the capital equipment that places and cures the composite material, and to meet aero-smoothness requirements, it would be necessary to build the composite laminate on a male lay-up mandrel, as a complete barrel section, and cut it into sections and transfer them to female cure tools. Laying up the skin as a barrel on a mandrel would allow the equipment to operate at significantly higher rates, thus reducing the need for additional machinery, reducing capital requirements.